Composite materials, such as graphite-epoxy, have been used in the manufacture of aircraft structures for many years because they have high strength-to-weight ratios. Initially, only lightly loaded or non-critical parts were manufactured from composite materials. As the technology advanced, however, a greater percentage of aircraft parts, including critical structural members, were manufactured from composite materials.
In contrast to metallic materials which are typically joined using fasteners such as bolts or rivets, composite materials are typically joined using adhesives. As a result, elaborate tooling is often required during the assembly of composite parts to hold the parts in their respective positions while the adhesive cures. The costs associated with this tooling can be substantial for large aircraft structures that must be accurately positioned.
FIG. 1 illustrates a prior art joining apparatus 110 for joining a wing rib panel 102 to a wing skin panel 101. The joining apparatus 110 includes a base 112 and first and second flanges 114 and 116, respectively, projecting outwardly from the base 112 to form a gap 115. The wing rib panel 102 has a thickness 117 and is positioned in the gap 115 between the first and second flanges 114 and 116. The wing rib panel 102 is bonded to the joining apparatus 110 with a generous amount of adhesive 118. The wing skin panel 101 is positioned adjacent to the outer surface of the base 112, and is bonded to the joining apparatus 110 with a layer of adhesive.
As shown in FIG. 1, the gap 115 is significantly wider than the wing rib thickness 117 to provide a considerable clearance between the wing rib panel 102 and the joining apparatus 110. This clearance is provided for a number of reasons. One reason is to prevent the adhesive 118 from being scraped off of the sides of the wing rib panel 102 as the wing rib panel 102 is inserted into the gap 115 during assembly. Another reason is to provide a path for excess adhesive 118 to escape as the wing rib panel 102 is inserted into the gap 115. Without this escape path, it may be difficult to fully insert the wing rib panel 102 into the gap 115 when the gap 115 is filled with the adhesive 118. Because of the large clearance between the wing rib panel 102 and the gap 115, tooling (not shown) is generally required to hold the wing rib panel 102 in the correct position relative to the wing skin panel 101 while the adhesive 118 cures.
There are a number of shortcomings associated with the joining apparatus 110 described above. For example, one shortcoming is the complex tooling that may be required when the joining apparatus 110 is utilized on large structures, such as aircraft wings having multiple wing rib panels 102 bonded to the wing skin panel 101. Such tooling may include, for example, multiple sub-fixtures, each configured to hold an individual wing rib panel 102 in its correct position relative to the wing skin panel 101. As mentioned above, the costs associated with this tooling can be substantial.
Another shortcoming of the joining apparatus 110 is the lack of bond pressure exerted on the adhesive 118 between the wing rib panel 102 and the first and second flanges 114 and 116 during curing, as a result of the oversized gap 115. Bond pressure can be a significant factor in developing sufficient bond-line strength. Accordingly, the bond-line that develops between the wing rib panel 102 and the joining apparatus 110 may not be as strong as it could be had sufficient bond pressure been applied. As a result, a design utilizing the joining apparatus 110 may have to provide more bond area than would otherwise be required to adequately join the wing rib panel 102 to the wing skin panel 101. This additional bond area can increase airframe weight and/or decrease aircraft performance.